Annular combustor for gas turbine

ABSTRACT

A combustor for producing a combustion gas stream turns the gas exiting therefrom to direct the gas stream at a first stage nozzle at a more advantageous angle of attack. The combustor is formed of a plurality of sections, each section containing a plurality of U-shaped steam coils placed askew to the linear axis if the combustor, the bend of the tubes forming the dome and the legs forming the respective inner and outer surfaces of the annulus.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a Continuation-in-part of U.S. application Ser. No.274,660, filed June 17, 1981, now U.S. Pat. No. 4,384,452 which is adivision of U.S. application Ser. No. 47,571, filed June 11, 1979, nowU.S. Pat. No. 4,314,442.

This application is also a Continuation-in-part of U.S. application Ser.No. 224,496, filed Jan. 13, 1981, now U.S. Pat. No. 4,438,625 which is adivision of U.S. application Ser. No. 954,832, filed Oct. 26, 1978, nowU.S. Pat. No. 4,272,953.

FIELD OF THE INVENTION

This invention relates to improvements in the reheat gas turbine and,more specifically, to a reheat gas turbine/steam turbine combined cyclewith steam cooling of the first stage nozzle vanes contained in both thegas generator and power turbine. The invention further relates to aprocess of performing work using a steam cooled reheat gas turbine. Thisinvention is equally applicable to simple cycle gas turbines.

DESCRIPTION OF PRIOR ART

In U.S. Pat. No. 4,272,953, applicant has disclosed that secondgeneration, high-cycle pressure-ratio, high-firing temperature gasgenerators can be used in the reheat gas turbine/steam turbine combinedcycle to yield increased efficiency and output heretofore unexpectedfrom reheat gas combined cycles. A novel reheat gas turbine withoutintercooling combined with a steam turbine is further disclosed inapplicant's pending application, U.S. Ser. No. 224,496, filed Jan. 10,1981. In this pending application, the reheat gas turbine comprises ajuxtaposed and axially aligned gas generator and power turbine in whichgas flow through the gas generator, reheat combustor and power turbineis substantially linear throughout.

In U.S. Pat. No. 4,314,442, applicant has disclosed that steam extractedfrom the steam turbine in a reheat gas turbine/steam turbine combinedcycle can be used to cool the reheat gas turbine vanes and rotatingblades. Steam cooling is shown to offer advantages over cooling with airor water at the higher temperatures which are possible in the reheat gasturbine/steam turbine combined cycle described in U.S. Pat. No.4,272,953. As described in U.S. Pat. No. 4,314,442, the extracted steamis directed as an external thermal barrier over the reheat gas turbinevanes and rotating discs and blades from internal steam plenums withinthe respective vanes and blades. Internal steam convective cooling ofthe gas turbine rotating blades is also beneficially used. The heatedcoolant steam is eventually ejected into the gas stream. It has beenfound that gas generator work is obtained by ejecting the coolant intothe gas stream and reheat pressure and reheat temperature are raised aswell.

In pending applications, U.S. Ser. Nos. 416,171; 416,172; 416,173 and416,275; all filed Sept. 9, 1982, applicant has disclosed that the gasgenerator and power turbine casings, struts, inner barrels, exhaust hoodand discs can be cooled by steam to allow a higher reheat firingtemperature. These parts are cooled in a computer-controlled mannerwherein the tip clearances of the compressor and turbine blades can bereduced and minimized under all operating conditions to improve cycleoutput and efficiency. A closed steam cooling system is disclosedwherein the coolant steam after being heated is re-introduced to thesteam turbine to form a second steam reheat which results in greatersteam turbine output at a higher combined cycle efficiency.

As shown by applicant, there are distinct advantages to utilize steam asa gas turbine blade coolant for high-inlet temperature levels of 2600°F. for both the simple and reheat gas trubines operating in the combinedcycle mode. For example, steam is twice as effective a coolant as airand is not as harsh or as difficult to control as water. All ofapplicant's U.S. patents and pending applications as described above areherein incorporated by reference to illustrate the advantages of steamas a blade coolant.

Air has been extensively used as a blade coolant. The emphasis inapplying air as a coolant has been to accomplish the most effectivecooling possible and thus allow raising the turbine inlet temperature.The air coolant has been given only secondary consideration as a workingfluid. However, aircraft engine manufacturers are presently usingthermal barrier ceramic coatings for the blading whereby less air isconsumed for cooling, leaving more air to do full work.

On the other hand, when considering steam as a coolant, as inapplicant's prior patents and pending applications, emphasis can also beplaced on the steam being a working fluid. Pressure drop is not ascritical since water can be pumped to 3000 psia pressure for only 10BTU/lb. The old criteria of cooling effectiveness does not hold the samemeaning where it is now possible to heat the majority of the steam tothe maximum while at the same time provide thermal shielding of theblading through an exterior laminar sublayer of steam which comprisesonly a small amount of the total steam flow.

An integrated propulsion nozzle is being applied to the "Rolls-RoyceRB-211-535 E4" engine for the "Boeing 757" aircraft. Performance gainsgiving 2% lower specific fuel consumption are realized where both thefan and the hot-core flows exhaust through a single propelling nozzle.High supersonic drag over the afterbody is reduced to a lower subsonicdrag. The "RB-211" exit nozzle measures several feet in diameter.However, the same general concept can be applied to very small steamnozzles measuring ten to twenty thousandths of an inch whereby steamwith its much lower viscosity, higher specific heat, and higher Mach 1velocity than air can be substituted in principle for the fan bypassair.

Historically, gas turbine first stage nozzle vanes have been subject toleading edge burning and low cycle fatigue cracking and the trailingedge region has been prone to develop hot spots, cracks and burn areas,particularly on the suction sides. Shower-head weep holes at the leadingedge of the nozzle vanes have been added to provide cooling thereof, butthe very thin trailing edge portion has been difficult to properly cooldue to lack of adequate space and available thickness.

Various methods of air cooling both stationary vanes and rotating bladeshave been patented and are in use today in aircraft and industrial gasturbines. U.S. Pat. No. 3,628,880 granted Dec. 21, 1971 to Robert J.Smuland et al relates to an air cooled vane utilizing impingementcooling, convection cooling with turbulence promoters and external filmcooling. U.S. Pat. No. 3,628,885, granted Dec. 21, 1971 to James E.Sidenstick et al relates to an air cooled rotating blade utilizing aserpentine flow path and turbulence promoters as well as impingementcooling and film cooling. U.S. Pat. No. 4,153,386 granted May 8, 1979 toJohn A. Leogrande et al relates to vane leading edge air cooling andpressure requirements thereto. These U.S. patents, however, do not applysteam as the coolant nor do they relate to an integrated nozzleconfiguration whereby the coolant steam is also used efficiently as aworking fluid in a binary steam/gas system as in the present inventionwhich is more fully described below.

SUMMARY OF THE INVENTION

In accordance with the present invention, improvements in gas turbinefirst stage nozzle vane efficiency are achieved by the integration ofmultiple steam nozzles with the first-stage gas nozzle vanes to form agas/steam binary flow system. Steam is first used as a vane coolant andsubsequently expanded and accelerated for work extraction by therotating blades as the hot gases pass around the nozzle vanes. A firststage 360° integrated critical flow exit area controls the mass flow andinlet pressures of both the steam and the hot gas. The present inventionis applicable to both simple and reheat gas turbines with or without acombined steam turbine.

Additionally, the present invention provides improved fluid cooling ofthe nozzle vane by altering the airfoil shape to form an integratedgas/steam nozzle vane with improved exterior steam thermal barrier filmshielding. By aligning the combustor to serve as a gas turning device toreduce the angle of attack of the gas to the individual vanes, the gasturning requirement of the first stage nozzle can be reduced. The shapeof the individual vane is altered by reducing the radius of curvature ofthe surfaces. Such reduced curvatures enhance steam thermal barrier filmshielding since a more stable laminar sublayer portion of the exteriorsteam boundary layer can be formed.

This invention presents an approach for integrating a steam nozzle withthe gas nozzle whereby nozzle vane cooling by internal steam flow isfirst accomplished. Serpentine reverse-flow cooling of the vane body isprovided for internal vane body cooling. A novel concept for internaltrailing edge "shock wave cooling" is set forth whereby the coolantsteam laminar sub-layer within an internal trailing edge nozzle isdestroyed to greatly improve heat transfer in the vital trailing edgeregion. Alternately, for sonic nozzle steam flow at the trailing edge,steam diffusion separation from sonic to subsonic velocity can likewisebe provoked to destroy the sub-layer and enhance trailing edge heattransfer.

Shock wave cooling as in the present invention can be described asfollows: Airflow (gas or steam) across a shock wave always decreasesfrom supersonic to subsonic velocity. The strength of the shock wavevaries with the amount of the deceleration of flow velocity. A strongshock wave absorbs energy. In a properly designed convergent/divergentnozzle, the flow downstream of the throat is supersonic. An internalsteam nozzle located at the vane trailing edge can be designed to act inthis fashion. An induced shock wave formed by over-expansion downstreamof the throat can be weak or strong depending on the degree ofdeceleration. As the shock wave grows stronger, the thin sluggish layerof flow (steam or gas) near the surface known as the boundary layer isinfluenced by the shock wave. The boundary layer attempts to reverseitself aft of the shock and flow upstream towards the shock, causing aphenomenon known as "shock-induced boundary layer separation". Aseparated boundary layer creates a turbulent wake aft of the shock. Thisreversal flow destroys the laminar sub-layer insulation next to thesurface and exposes the bare metal to the coolant. A much greater heattransfer coefficient results for much greater cooling rates.

In an opposite sense, controlling and limiting the velocity to Mach 1 onthe exterior aft suction side of the vane will prevent a shock wavedeveloping and prevent boundary layer separation of the exterior steamthermal barrier. A laminar exterior steam film can be maintained toreduce drag and to provide steam film insulation with lower heattransfer rates. Airflow across a shock wave is clearly explained in U.S.Pat. No. 3,952,971 to Richard T. Whitcomb, granted Apr. 27, 1976 forAirfoil Shape For Flight At Subsonic Speeds. The intent given in thisreferenced patent is to reduce drag and increase lift and this airfoilis presently being applied to the "Harrier" vertical take-off andlanding military aircraft and the "Boeing 757 and 767" commercialpassenger aircrafts. The same aerodynamic principles are applied in thepresent invention to reduce or increase, as the case may be, heattransfer primarily and secondarily to reduce overall drag. Suchprinciples have never been applied or suggested in gas turbine vanesbefore for improved internal cooling and external reduced heat transfer.

A further objective of this invention is to set forth a method ofimproving the efficiency of the integrated nozzle through reducedsurface friction losses and reduced trailing edge losses by reducing thetrailing edge wedge angle from the customary ten degrees to zero. Vectormomentum is thus conserved.

A still further objective of this invention is to provide a segmentedsteam or low BTU fuel gas cooled annular combustor liner whereby thecombustor can be assembled and disassembled in individual segmentsaround a continuous shaft and whereby the joints between the individualsegments can be locked together to form a complete annular combustorliner without any horizontal joints of flanges which cause thermalstresses and uneven thermal expansion. Heretofore, attempts to providean annular combustor liner that can be fit around a continuous shafthave not been successful for single shaft industrial gas turbines eventhough the superior characteristics of even temperature distribution tothe first stage nozzle of the annular combustor are well known.

Still another objective of this invention is to provide a combustorliner with approximately 25 percent greater flow length for a givenaxial length by changing the axial direction of the gas, thus, makingthe gas flow askew to the axis of the machine and causing the gas toexpand radially outwardly as the gas is heated by the burning fuel.Cooling air passes around and through the combustor liner to mix withthe near stoichiometric primary flame to temper the mixture to an evensculptured temperature profile to the first stage nozzle. The individualhigh temperature spikes of the individual cylinder combustors of the"can" type combustor are eliminated, such spikes causing undesirablepulsating forces on the downstream rotating blades and possible failuresthereof. The temperatures of the gas near the nozzle inner and outersidewalls can be reduced to a lower temperature by some 300° F. with themaximum temperature being located approximately two-thirds outwardlyfrom the inner root of the vanes and rotating blades. Such profilingwith annular combustors is well known by those familiar with the art ofgas turbines.

The mentioned objectives of the present invention and other objectiveswill become apparent upon a more detailed description of the integratedgas/steam nozzle and the integrated combustor/first stage nozzle and canbe ascertained by the detailed description of the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a graph of profile loss coefficient versus exit Mach numberfor typical convergent/divergent and convergent gas turbine nozzles withnormal gas losses shown by the solid lines and losses associated withthe integrated gas/steam nozzle of the present invention shown by thedashed lines.

FIG. 2 is a graph of sonic velocity versus molecular weight of severalgases at various temperatures.

FIG. 3 is a perspective view in partial section of the vane assembly ofthis invention.

FIG. 4 is an enlarged cross-sectional view of the vane assembly takenalong line 4--4 of FIG. 3.

FIG. 5 is an enlarged partial cross-sectional view of the vane bodyshowing impingement and body serpentine coolant steam flow paths.

FIG. 6 is an enlarged partial cross-sectional view of the vane bodyshowing a trailing edge convergent/divergent coolant steam nozzle.

FIG. 7 is an enlarged partial cross-sectional view of the vane bodyshowing an alternate configuration of a trailing edge convergent coolantsteam nozzle.

FIG. 8 is an enlarged partial cross-sectional view of the vane bodyshowing the trailing edge steam nozzle and trailing edge coolantpassages between nozzles.

FIG. 9 is a graph of surface Mach number versus surface distancedownstream from the vane leading edge.

FIG. 10 is a profile of a sonic shock wave showing the destruction ofthe coolant steam sub-layer.

FIGS. 11 and 11A is a perspective cut-a-way view of the segmentedannular combustor liner used to accomplish gas turning.

FIG. 12 is a partial vertical cross section of FIG. 11 illustrating airand gas flow to and through in the combustor liner of FIG. 11.

FIGS. 13a and 13b illustrate the cross-sectional shape comparison ofalternate vanes with different angles of attack.

FIG. 14 is a profile of a sonic shock wave emanating from the sheardisturbance in the coolant steam boundary layer.

FIG. 15 is a profile showing the throat width and trailing edge shockwave at the throat of the integrated gas/steam nozzle.

DETAILED DESCRIPTION OF THE INVENTION

Like reference numbers will be used to identify like parts in thefollowing description of the preferred embodiments.

A typical profile loss ccefficient graph for both convergent andconvergent/divergent gas turbine nozzles is presented in FIG. 1 whereprofile loss coefficient is plotted against exit Mach number. Theprofile loss coefficient for any given Mach number reflects the loss intotal pressure taking place and specifically the loss in potential exitvelocity. Such loss is reflected in less mechanical work extraction bythe rotating blades. This graph shows clearly that the profile loss fora Mach 1 (convergent) nozzle is low, being about 2 percent, whereas aconvergent/divergent design has much higher losses, ranging from 5 to 7percent. Superimposed on the graph of FIG. 1 are dashed lines showingloss savings achieved using the integrated gas/steam nozzle of thisinvention as more fully described below. Surface frictional loss islowered due to blanketing steam having one-half the viscosity of thegas. Also, the trailing edge exit loss can be reduced by eliminating thecustomary ten degree wedge angle. The wedge angle is the angle formed atthe trailing edge by the trailing edge suction side surface and thetrailing edge pressure side surface, the lines forming the wedge anglebeing tangent to the curvature at the trailing edge extremities. Mach1.1 to 1.2 efficiency savings are of particular interest.

When considering steam as a blanketing fluid for film cooling(shielding) and as a working fluid, surface sonic velocity becomesimportant. Steam has a notably higher sonic velocity than air (gasstream) for any given temperature. For instance, at 800° F. the Mach 1velocity is 23.0 percent higher for steam and at 2600° F. this velocityis 21.9 percent higher. A graph of sonic velocity versus gas molecularweight is given in FIG. 2. Note that air at 2000°0 R. (1540° F.) has asonic velocity of about 2100 ft/sec whereas steam at this temperaturehas a Mach 1 velocity of 2700 ft/sec. Also note that hydrogen has a veryhigh sonic velocity of some 8400 ft/sec. for this temperature. A mixtureof 96.4 percent steam and 3.6 percent hydrogen by weight would have asonic velocity of about 3100 ft/sec. Likewise a mixture of steam and air(50 percent by volume, 38 percent by weight) would have a sonic velocityof about 2400 ft/sec.

Hydrogen is known for its excellent cooling characteristics. Thus, inaccordance with the present invention, steam as a coolant can be"sweetened" with hydrogen for the reheat first-stage nozzle if suchaddition results in a mixture that is well below the flammability limitsof 4% by volume in air. Most of the hydrogen added can be oxidized inthe reheat combustor, at least the portion that enters the primary heatrelease zone. Although, "spiking" steam with hydrogen offers interestingpossibilities, hydrogen is most difficult to handle and store. Likewise,"spiking" cooling air with steam has distinct advantages and willimprove the cooling capabilities of pure air in accordance with themixture's sonic velocity.

There is another important consideration when applying steam blanketingfor nozzle vane cooling and as a working fluid and that consideration iscooling effectiveness decay due to the mixing and diffusion of the steamand gas downstream of the steam nozzles which are contained in theleading edge of the integrated nozzle vanes of this invention. The rapidtemperature rise of the majority of the coolant steam increases the workcapability but decreases the steam cooling ability by also increasingthe film temperature. Both aspects must be considered. This decay incooling effectiveness points to the need for additional downstreaminternal vane cooling and also indicates that such cooling should be ina reverse direction from the vane trailing edge upstream to the leadingedge steam nozzles. It is important to maintain blade metal temperaturesas even as possible to avoid adverse thermal gradients and unevenexpansion which is a cause for low cycle fatigue problems.

Gas turbine nozzles provide four basic functions which are vital andnecessary to convert thermal energy to mechanical work. These functionsare listed as follows:

1. Accelerate the gas,

2. Change and control gas direction,

3. Control gas-mass flow,

4. Reduce gas temperature to rotating blades.

The gas must be accelerated efficiently to a given design velocitythrough a pressure drop process. The first stage nozzle is unlikedownstream rotating blading or subsequent nozzles because its approachvelocity is rather low--being about Mach 0.1. Greater acceleration isrequired. Work after acceleration can then be extracted from the gas byapplying impulse or a combination of impulse and reaction rotatingblading. Secondly, the gas must strike the rotating blades at the rightangle for maximum work output and extraction efficiency. The gas turningangle of the first stage nozzle is generally about 70 to 72 degrees withrespect to the centerline of the turbine. Thirdly, the mass flow must becontrolled. This third function is generally accomplished by designing acritical area first stage nozzle where sonic (choking) velocity preventsany further increase in mass flow for any given inlet pressureregardless of the outlet pressure. The fourth function is achieved by asignificant temperature drop of the gas when pressure drops and the gasaccelerates.

Blade cooling using air as the coolant introduces a parasitic loss andthe cooling air has no basic function of providing additional work.Extra work comes from the increase in firing temperature.

When considering an integrated binary gas/steam nozzle, there are fouradditional functions besides the four basic functions previously listedthat must be considered when the steam is used not only as a coolant,but also as a working fluid. The additional four functions are asfollows:

5. Reduce friction loss,

6. Mix cool steam with hot gas,

7. Provide film cooling (shielding),

8. Increase exit velocity past Mach 1 without extra loss.

Item 5 deals with surface friction. Steam at a lower temperature andlower basic viscosity has half the viscosity of cooling air and,therefore, if steam is introduced efficiently and effectively near thevane leading edge, then less overall frictional loss will occur acrossthe downstream surfaces.

The majority of the steam injected into the gas stream from the steamnozzles in the vane, Item 6, must be heated by the hot gas as early aspossible to obtain maximum steam work. The heating is accomplishedthrough diffusion and mixing in the outer regions of the boundary layer.

The nozzle vane must also provide the function of establishing a stablelaminar flow sub-layer over most of the outer surface whereby aninsulative shield is formed to reduce heat transfer, Item 7.

The difference in Mach 1 velocities, Item 8, of steam versus air (gas)gives rise to the distinct possibility of being able to accelerate themain-stream gas to a Mach value of perhaps 1.15 without suffering theusual high trailing edge shock wave loss.

NOZZLE VANE CONFIGURATIONS

Typically, there are 24 to 48 vanes in the 360° entrance nozzle for a300 lb/sec. airflow turbine. FIG. 3 shows a typical nozzle segment 20containing one vane 22. With reference to FIGS. 3 and 4, nozzle vane 22is a "fat-bodied" type nozzle vane airfoil. Nozzle vane 22 comprises aremovable, leading edge portion 24 connected to a source of steam whichis fed through plenum 26. Leading edge portion 24 has a ceramic thermalbarrier outer coating 28 and contains "shower-head" weep holes 30 atnose 32. Weep holes 30 curve back toward trailing edge portion 34 ofvane 22. Two leading edge steam nozzles 1 and 2 are contained in leadingedge portion 24. Vertically spaced fins 40 of approximately 0.010 of aninch thick connect forward shell 42 with plenum core 44 and form nozzlechannels 46. Fins 40 provide flexibility for differential expansionbetween plenum core 44 and forward shell 42 and also provide extendedheat transfer surface. Leading edge portion 24 can be either cast ormade up of thin wafers fusion-bonded together, a new technique currentlybeing developed. Leading edge portion 24 is positioned on vane body 48by keyway 50. Vane body 48 formed by vane walls 52 and 54 is cast aspart of inner and outer side walls 56 and 58, respectively, FIG. 3. Itshould be noted that side walls 56 and 58 can be shaped to provideproper side wall acceleration and gas stream flow. There can be one ortwo vanes 22 per vane segment 20.

The cooling of leading edge portion 24 with steam can be understood byreferring to FIG. 4. Coolant steam entering plenum 26 is directed toleading edge steam plenum 60. From plenum 60, the coolant steam passesthrough nozzle 62, is split by leading edge flow divider 64 and directedto nozzle channels 46 of respective leading edge steam nozzles 1 and 2.A small portion of the steam passing through nozzle 62 enters a swirlchamber 66 through tangential flow holes 68. Swirl chamber 66 iscontained in nose 32. Steam then weeps from swirl chamber 66 to theextreme leading edge and exits leading edge 24 through weep holes 30 innose 32. The majority of the coolant steam is directed to nozzlechannels 46 which contain turbulator ribs 70. A small portion of steamfrom nozzle channels 46 is weeped to the gas stream through weep holes72 as shown in FIG. 4.

FIGS. 4 and 5 also illustrate the manner in which vane body 48 is cooledwith steam. A separate external supply of steam enters dual "pant-leg"ducts 74 and 76 from entrance plenum 78 at one common connection atouter sidewall 58 for cooling vane body 48. Ducts 74 and 76 are sealedat the inner end and are free to float to allow for differential thermalexpansion. One edge 80 of inner duct walls 82 and 84 of respective ducts74 and 76 runs in a groove in trailing edge nozzle block 86 as shown inFIG. 6 and is not welded or bonded to body 48. Coolant steam enteringsteam plenum 78, FIG. 3 and discharged to the forward portion of ducts74 and 76 first travels towards trailing edge 34 where the steam isaccelerated to impinge against the interior of vane walls 52 and 54 neartrailing edge nozzle block 86. The impingement shown by arrows 87 and 89in FIG. 5 provides greatly increased heat transfer in a critical area.The coolant steam then takes a serpentine route back and forth betweeninner and outer side walls 56 and 58, respectively, through wall ducts88 adjacent to each of vane walls 52 and 54 and back-flows to steamnozzle 3 and central steam plenum 90. Serpentine flow is common practicefor air cooling. The vertical dams 92 and serpentine flow passages 88are better seen in FIG. 5. The serpentine ducts 88 have thin crossribs(cleats) 94 to promote turbulence. A 30 to 40 psi pressure drop isestimated to take place between ducts 74, 76 and plenum 90. Steampassages 96 near trailing edge 34 allow coolant steam from flow passages88 to flow through cooling channels 98 and 100 in inner and outer sidewalls 56 and 58, respectively. Steam from central steam plenum 90 flowsagain parallel to gas flow to trailing edge steam nozzle 4.

The "pant-leg" ducts 74 and 76 are made of thin sheets of stainlesssteel material approximately 0.010 of an inch thick which form inner andouter duct walls 82, 83, and 84, 85 of the respective ducts.Alternately, the duct walls can be laminated from two layers of 0.006 ofan inch thick material with the inner laminate being fabricated withdimples 102 (see FIG. 3) or upsets 0.003 of an inch deep on a 1/4 inchgrid. Dimples 102 separate the laminates. The dead space betweenlaminates will provide insulation and reduce heat transfer between steampassing through ducts 74, 76 and steam traversing the serpentine flow.The added pressure inside each duct 74 and 76 forces respective outerwall ducts 83, 85 against the larger vertical dams 92 to form areasonable seal with low leakage. There are hollowed out bridges 104between ducts 74 and 76 to prevent ballooning due to the differentialpressure. The reaction forces are taken against hollow-end supports 106.

The vane suction side vane wall 54 from steam nozzle 3 all the way totrailing edge 34 is perfectly flat to encourage a stable exteriorlaminar sub-layer of steam from nozzle 3. The main steam velocity is atMach 1 where steam exits nozzle 3. The pressure side vane wall 52 ofvane 22 between leading edge steam nozzle 2 and trailing edge steamnozzle 4 has only a very slight curvature to help maintain the laminarsub-layer steam film. The main stream is being continuously acceleratedand therefore the boundary layer remains stable. The trailing edgepressure surface 108 takes a greater curvature at nozzle 4 so thatpressure surface 108 becomes parallel with trailing edge suction sidesurface 110 thus eliminating the normal wedge or convergent angle. Steamfrom nozzle 4 fills the downstream void as will be subsequentlydiscussed.

Leading edge steam nozzles 1 and 2 are subsonic. Nozzle 3 is sonic andnozzle 4 can be sonic or transonic as will also be explained. Thethickening of the trailing edge at 112 allows space for the "pant-leg"insert and the backflow serpentine arrangement as well as provides forparallel flow.

Reference is made to FIG. 6 which illustrates the trailing edgeconvergent/divergent transonic cooling steam nozzle 4. FIG. 7 shows theMach 1 alternate design. The rather thick trailing edge 34 (0.090 to0.150 of an inch for a typical machine having a 300 lb/sec air flow)made possible by steam nozzle 4 gives rise to a different manufacturingapproach not possible with the conventional thin-edge designincorporating low-flow weep holes or low velocity gill escape slots. Thethicker trailing edge allows the nozzle vane walls 52 and 54 andsidewalls 56 and 58 to be cast integrally in either single or dual vanesegments. Trailing edge walls 108 and 110 can be cast thick enough toinsure integrity. It is well known that it is difficult to cast thinsections of 0.020 of an inch or thereabout with integrity even though"HIPPING" might help. Alternately, trailing edge 34 can be cast solid.Subsequently, a slot can be precisely machined into the trailing edge toprovide an accurate fit for trailing edge nozzle block 86.

Trailing edge nozzle 4 has rather small and intricately shaped passageswhich require precision machining and fabrication. The nozzle block 86,for example, can be made in two halves, precisely machined and thenfusion-bonded together. Nozzle block 86 can then be finish-machined andinserted into the trailing edge slot. Vane body 48 and nozzle block 86can be fusion-bonded together to form one piece. Various materials canbe selected to form nozzle block 86 and vane body 48. Nozzle block 86can be made out of a material suitable for maximum erosion resistanceand body 48 can be cast out of a material having good low-cycle fatigue,corrosion, high-temperature resistance and castability properties.

Referring to FIGS. 6, 7 and 8, a small portion of steam from duct 74(pressure side) is fed to the space 114 between trailing edge nozzles 4which are vertically spaced within trailing edge 34. The steam impingeson the extreme trailing edge 116 of vane 22 for cooling beforebackflowing to the nozzle 4 entrance. Passageways 118 from duct 74provide the coolant steam flow. Inserts 120 are placed in drilled holes122 in nozzle block 86 to force the steam to flow at a high velocitynear the sidewall surfaces. The inserts can have broad screw threads 124to give further steam path travel.

Vane Internal Cladding

It is important that the vane body metal temperature be as uniform aspossible to prevent adverse thermal gradients. In addition to thetrailing edge cooling and the serpentine backflow arrangement, a furtherapproach is proposed to help even out the temperature from the trailingedge to the mid section.

A thin cladding 126 (about 0.030 of an inch thick) of a material havinga very high heat conductivity may be fused to the internal surfaces ofvane walls 52 and 54. Copper or silver can be used. Silver has a meltingpoint of 1760° F. and a conductivity of 241 BTU/h/ft² /°F./ft and copperhas a melting point of 1980° F. and a conductivity of 227 BTU/h/ft²/°F./ft whereas superalloys have conductivities ranging from 8 to 14BTU/h/ft² /°F./ft. Copper or silver would be twenty times better as aheat conductor. Aluminum, although a good heat conductor, has too low amelting point to be considered (1220° F.). Copper would probably be thebest choice because of its higher melting point.

Techniques of fusing copper to high temperature superalloys have beendeveloped by the Department of Energy through the water-cooled bladeresearch program and can be applied to the steam cooled vane.

The high conductivity metal cladding 126 will conduct heat from the thintrailing edge walls to the thicker body portion so that the vane portionnear the alignment keyway 50 will be at an estimated temperature ofabout 1100° F. whereas the trailing edge metal will be at an estimatedtemperature of about 1300° F.

Steam/Gas Nozzle Integration

The integration of the multiple steam nozzles with the gas nozzle willnow be discussed. The steam must cool the vanes but at the same timeexit to the gas stream at a velocity close to the gas-stream velocityand parallel with the flow so that the boundary layer will not bedisrupted. The conventional weep holes at their 30 degrees angle in aircooled blades will allow the coolant to blow directly through theboundary layer if the velocity is too great and cause greater heattransfer instead of less. The steam flow exiting tangential to the vanesurface avoids this difficulty for off-design operation. Mach numbersmust be considered as well as the various heat-transfer aspectsinvolved.

An example of the operation of the integrated nozzle of the presentinvention can be described for a reheat gas turbine having a 40 cyclepressure ratio and a nominal turbine inlet temperature of 2600° F. Gasat 2600° F. and 564.1 psia allowing a 4% combustor pressure drop expandsto the critical Mach 1 exit pressure of 298 psia and 2190° F. at nearconstant entropy.

Reference is now made to FIG. 9 which has been derived for the reheatgas turbine operating at the above-stated conditions. The solid linesrepresent the gas and the dashed lines represent the steam after itexits from nozzles 1, 2, 3, and 4. It should be noted that even thoughthe gas stream and steam velocities are the same, the Mach numbers aredifferent because steam has a higher Mach velocity and the difference intemperature. The suction-side Mach velocity rises rapidly to nozzle 3about 50% of distance of the surface where Mach 1 choking takes place.The gas continues at Mach 1 to nozzle 4. The pressure side surface gasMach number remains rather low for fifty percent of the path and thenturns upward to nozzle 4, Mach 1. The greatest increase is near the end.Suction-side steam enters from nozzle number 1, at its Mach number ofabout 0.55 and progresses subsonically to point 4 where its Mach numberis about 0.9. Likewise, pressure-side steam enters from nozzle 2 at itsMach 0.15 velocity and flows subsonically to nozzle 4, Mach 0.9. Thesteam passing through nozzle passages 46 takes a non-recoverablepressure drop of perhaps 15 psi and expands subsonically throughpassages 46 to exit points 1 and 2 of FIG. 9.

Suction side steam nozzle 1 steam flow explanation follows for thepreviously stated operating conditions. The steam at a predeterminedpressure of 625 psia and 645° F. expands and exits to the gas stream ata velocity of 1000 ft/sec. at nozzle 1, FIG. 9. The steam then starts toheat as it expands to the trailing edge nozzle 4, taking a somewhat "S"curve due to the heat being progressively added. The final trailing edgeexit velocity can be slightly higher than the main gas stream due to thesteam's higher enthalpy (driving force), velocity being a function ofthe square root of the enthalpy drop.

The pressure side steam exiting nozzle 2, FIG. 9 is at a lower exitvelocity. Its velocity is about 300 ft/sec. Flow ribs restrict the flowby reducing the pressure before the exit. An exit velocity of about 300ft/sec. corresponds to the Mach number shown in FIG. 9. Note that theleading edge steam remains subsonic all the way over the vane bodysurface and approaches Mach 0.9 at the trailing edge. If the steamexited at 1800° F., the Mach 1 velocity would be about 2750 ft/sec

Steam pressure in leading edge steam plenum 60 is not critical and canvary somewhat. The designer has freedom in selecting the design pressureand temperature to fit the cold reheat pressure or an extraction pointon the steam turbine. Further, if desuperheated steam (saturated) isused, the steam will become supersaturated as it expands through thefirst nozzle 62 to divider 64. The specific heat can rise to a valuegreater than one and heat transfer will be improved.

In the example of the 40 cycle pressure ratio reheat gas turbine, steamat a pressure range of 580 to 780 psia, saturated (481° to 515° F.) orwith 10° to 20° F. superheat, flows from dual plenums 74 and 76, FIG. 4,to central steam plenum 90. The steam impinges on the area near steamnozzle 4 entrance and makes a 180 degree reversal. It then backflowsthrough serpentine passages to plenum 90 as described previously. Thesteam, having lost approximately 30 psi pressure and having been heatedin the process to a temperature of 800° to 1000° F. is then ready toexpand through nozzles 3 and 4, FIG. 4. Approximately 50 percent of thesteam accelerates to Mach 1 at each nozzle throat. The steam, in thecase of nozzle 3, discharges to the main stream at Mach 1 velocity. Thethroat exit velocity is about 2000 ft/sec. The steam immediately startsto heat and expand at Mach 1 velocity, approximately 2350 ft/sec.

Nozzle 4 steam has two alternate flow possibilities, see FIGS. 6 and 7.First (FIG. 6), the steam nozzle can be designed to diverge from throat128 to allow supersonic flow and over-expansion. The nozzle duct 130after diverging has a constant cross-sectional area being altered onlyslightly to accomodate boundary layer thickness growth, heating frictionand trailing edge heat absorption. The steam is heated as it flows. Thenozzle again diverges at 132 near the exit, but the over-expansioncannot accomodate any further divergence and a shock wave is provoked.The shock wave is depicted at 134. The shock wave and downstreamreflections thereof destroy the boundary layer at the extreme trailingedge and the laminar sub-layer thermal barrier steam film formed in duct130 to expose the metal surface at the extreme trailing edge directly tothe steam. The surface heat of friction is mixed with the main stream.Heat transfer is thus enhanced.

Reference is made to FIG. 10. Shock waves are created by a disturbanceclose to the surface next to the laminar (viscous) sub-layer. A severenon-viscous shearing, tearing or ripping takes place that generates astrong sonic output. There is a "bursting" effect. The sound wavescannot escape and they build to form a pressure barrier or shock wave asis well known. It is the action of the flow over the surface or thesurface through the gas or steam that creates the shock wave and not themain body of the flow or gas region.

The second alternative (FIG. 7) is for trailing edge steam nozzle 4 tobe designed so that the steam will expand to sonic velocity and continueat sonic velocity as it is heated. The steam duct 136 cross-sectionalarea downstream from throat 138 is increased only in the amountnecessary to accomodate the change in density and temperature. The areachange can be calculated by the following formula derived from the sonicvelocity formula V_(s) =√KgRT and Boyles law PV=WRT. ##EQU1## where A₁is the original throat area, A₂ is the downstream area, T₂ is the higher(heated condition) temperature and T₁ is the upstream temperature. As anexample, if the steam is heated from 700° F. to 800° F., the area ratiochange would only be 1.042. This area would be increased slightly toaccomodate the thickening of the boundary layer. If the duct near theexit point is allowed to diverge rapidly at 140, FIG. 7, the steam willtry to diffuse. Separation will set in causing eddy current and reverseflow in the boundary layer. The boundary layer will be destroyed at anddownstream of location 140 and the heat of friction will be mixed withthe main stream. Heat transfer will increase.

A velocity loss takes place in either nozzle 4 configuration due to theshock wave or Mach 1 to subsonic velocity diffusion separation, but thesmall loss is considered acceptable in exchange for the improved heattransfer at the trailing edge where it is badly needed.

The gas stream passing trailing edge walls 108 and 110 run parallel witheach other as shown in FIGS. 6 and 7. Normally there is a ten degreewedge angle which causes a finite vector momentum loss. Velocity is avector quantity and the right angle components cancel each other. Themomentum can be expressed as follows:

    M(Ve-V.sub.o)=1/2MVe (1-cos θ/2)                     (2)

where M is the mass, Ve is the velocity of each side (assumed equal)before merging, V_(o) is the combined velocity after merging and θ isthe wedge angle. Note that only half of the velocity is effected by thewedge angle because the flow momentum is neutral at the center lineposition between vanes where there is no wedge angle.

The velocity loss (constant mass flow) for a ten degree angle calculatesto be about 0.2 percent. This value is small but worth saving whenconsidering that change in enthalpy (and work) varies as the square ofthe velocity and the 0.2 percent loss for velocity becomes a 0.4 percentloss for work.

The proposed trailing edge has a thickness of about 0.120 to 0.150 of aninch for FIG. 6 and 0.090 to 0.120 of an inch for FIG. 7 considering atypical 300 lb/sec air flow. The steam exiting out of nozzle 4 will beat a temperature of about 900° F. for an entrance temperature of 720° F.A temperature rise due to trailing edge heat transfer is calculated tobe about 100°0 F. The exit steam at 900° F. will mix with the 2200° F.gas stream. The steam will heat rapidly and double its volume (Boyle'slaw) to fill the partial void created by the lips of the nozzle andvertical spaces between each nozzle slot. The surface steam from nozzles1, 2 and 3 will also heat to fill the partial void. Steam is muchsuperior than air in filling the void because steam expands 1.6 timesgreater than air for a given temperature rise.

The steam nozzle area and slot width for nozzles 1-4 can be readilycalculated from mass flow rates knowing the steam nozzle exittemperature and specific volume.

Gas Turning Angle

Nozzle vanes must turn the gas rather sharply to direct the gas properlyto the rotating blades. The turning angle is generally 70 to 72 degrees.Curved vane surfaces are required and such curved surfaces enhance heattransfer from the gas stream to the vane. Recent innovations have beenmade whereby "blunt-nosed" wedge-shaped designs are being appliedwhereby the pressure and suction surfaces are much flatter. The surfacesare thus easier to film cool. Some aerodynamic loss occurs at theleading edge when applying this approach, but a low approach velocityminimizes these losses.

In the case of the steam cooled nozzle vane or any air or water cooleddesign, another method can be used. Reference is made to FIGS. 11 and 12which show a steam cooled annular combustor liner at 142. Combustorliner 142 can be tilted outwardly with a smaller entrance pitch linediameter than the outlet pitch line diameter. Compressed air with ageneral axial flow direction surrounds combustor liner 142 and thepressure is retained by outer casing 141. Fuel is fed to the inside ofliner 142 by fuel line 143. Compressed air enters combustor liner 142between adjacent liner coils 148 as indicated by arrows 147, FIGS. 11and 12. A set of cocked fuel nozzles 144, dome-turning vanes 146 andsteam combustor liner coils 148 which are askew can be applied to effectpart of the gas turning function. FIGS. 11 and 12 show one of multiplefuel nozzles 144, combustor dome 150 with turning vanes 146 and linercoils 148 arranged to provide 36 degrees gas turning which is half ofthe required gas turning angle of 72 degrees. The gas turning angle αplus the vane turning angle β thus generally will be equal. The turningvanes 146 in dome 150 comprise cooling coil tubing 152 covered withextended surface streamline sheathing 154. Combustor liner coils 148comprise cooling coil tubing 152 covered with an extended surfacesheathing 156. Steam can be passed through tubing 152 to providecombustor liner cooling. The heated steam can be diverted to a steamturbine, if desired.

The angle of attack on the nozzle vanes 22 is now reduced from 72degrees to 36 degrees and the leading edge shape can be substantiallyaltered. Also, the pressure side surface can be flattened so that thesteam from nozzle 2 can be more effectively administered to provide amore stable sub-layer laminar film. FIGS. 13a and 13b show,respectively, the prior art nozzle vane 23 and axial approach arrow 25and nozzle vane 22 for the 36° approach indicated by arrow 25' asachieved by the present invention. The pressure side gas will acceleratemore evenly from the stagnation point to the trailing edge which againhelps retain and stabilize the laminar sub-layer for thermal shielding.

A comparison of the two vane shapes is shown in FIGS. 13a and 13b. Notethe angle of attack, the changed nozzle vane leading edge shape and theflatter pressure side surface. It can be determined that the 36 degreeangle of attack vane has a leading edge suction surface with some 25percent reduced curvature. The 36 degree angle of attack vane also hasonly half the curvature radius of the 72 degree vane on the bodypressure side. These substantial reductions in curvatures provide lessaerodynamic drag and make possible surfaces which in turn encourage astable laminar sub-layer for better film shielding.

The turning of the air and gas by the fuel nozzles and the skew tubearrangement is shown in FIGS. 11 and 12. Steam tubes 152 with theirextended surface sheathing 156 are wrapped at a 36 degree angle withrespect to the turbine frame. Individual coils 148 are hairpin shaped toform the inner and outer annular combustor surface and are set at the 36degree angle. Combustor dome 150 is formed at the bend of the individualcoils 148. The tubes 152 can be of an inside diameter of one-quarter toone-half inch. The low velocity cooling air will flow from the exteriorof the annular liner 142 to the inside at the 36 degree angle where thevelocity will be much higher and at a 2 to 21/2 percent lower pressure.

Combustor dome 150 and the outer liner area contain openings 158 betweenthe individual hairpin shaped coils 148. These openings 158 arenecessary to allow the dual fuel nozzle 144 to be inserted inside theliner 142 and to provide an entrance for primary combustion air throughswirler 145. On each outer and inner surface of annular combustor 142are provided openings 160 near fuel nozzles 144 and between theindividual hairpin shaped tubes 152 to allow additional air to enter thecombustion zone for proper combustion and the desired fuel/air ratio forNO_(x) control.

Primary combustion air flows to the inside of liner 142 by means ofswirler 145 and turning vanes 146. Swirler 145 is configured to providea spiral spinning action for the fuel and air, the swirler flow patternarrows 149 and fuel flow pattern arrows 151 being shown in FIG. 12. Boththe swirler flow pattern 149 and fuel flow pattern 151 are askew to thelinear axis of the machine. Primary air passing through dome 150 isturned by vanes 146 to a flow path 153 also askew to the linear axis ofthe combustor liner.

Cooling combustion liners with compressor discharge air is becoming moredifficult where higher compressor discharge temperatures are encounteredat the higher compression ratios and where more air is needed to limitthe primary flame temperature to help control NO_(x) pollutants. Thesteam cooled liner of this invention makes available more air to controlNO_(x) because less air is needed to film cool the inside surfaces whichare subjected to radiant heat from the primary flame. Thus, the spacesbetween coils 148 can be reduced since the coolant steam passed throughthe coils removes a substantial portion of the radiant heat.

The superior design of the annular combustor is well known and attemptshave been made in the past to design an annular combustor that could besplit in half so that an annular combustor liner could be used on asingle shaft industrial gas turbine whereby it could be slipped over theshaft. To date no such arrangement has been successful due to thehorizontal liner flange and the thermal distortion created.

Aircraft gas turbines use modular construction with splined shaftingwhereby the non-split annular combustor liner can be inserted with theshaft removed. Then the splined shaft can be inserted inside the annularcombustor to mate the compressor shaft with the turbine shaft.

The proposed steam cooled combustor liner, with reference to FIG. 11,can be easily segmented into two, three, four or more parts. Theindividual segments can be slipped around a shaft and secured to theshell to form a single annular combustor. The combustor liner of FIG. 11shows three segments 162, 164 and 166. The ends of tubes 152 forming theouter surface of annular combustor liner 142 are associated with therespective connector flanges 163, 165 and 167 for each segment, theindividual flanges being connected at joint 168. Thermal expansion slits169 between the bolt holes and the outer flange edge allow fordifferential growth. The connector flanges are bolted to the turbinebulkhead. An annular steam plenum 180 can feed or receive steam fromtubes 152 forming the outer surface. The ends of tubes 152 forminginside combustor liner surface can be associated with a slip type jointcommonly used in gas turbines.

Combustor liner inside surface exit extremity 170 is fitted with themale portion of the slip joint. As separate annular steam plenum 182 canbe associated with tubes 152 forming the inner combustor liner surface.

There is no horizontal joint to contend with as the joint or jointsbetween segments 162, 164 and 166 run at a 36 degree spiral where theindividual hairpin tubes 152 join each other with reference being madeto FIG. 11a. Positioning clips 184 can be used whereby each segment willslip adjacent to the adjoining segment before being bolted to theturbine bulkhead. Male L-shaped extension 183 from one combustor linersegment or grouping of tubes 148 slip into female L-shaped opening 185of the adjacent combustor liner segment at segment joint 168 to positionand lock the adjacent segments together. No welding or bolting of thesegments is necessary. The individual segments can be screwed into thepositioning or locking clips. Clips at the dome area, midspan and thebulkhead are all that are necessary.

With reference to FIG. 12, it can be seen that the flame travel distancefrom fuel nozzle 144 to the first stage nozzle 18 will be longer withthe cocked fuel nozzle, inlet turning vanes and the skew steam coolingcoils. The flame travel or length will be about 24 percent longer for a36 degree turning angle where the hypotenuse of the vector triangleformed is equal to the reciprocal of the cosine of 36 degrees with theaxial flow length being equal to the cosine of 36 degrees. The added 24percent longer travel will make it possible to reduce the length of thecombustor about 15 percent, and still have added protection againstflame impingement on the first stage vanes, a distinct advantage up andbeyond film cooling advantages already mentioned.

Vane Exit Mach Velocity

As previously explained, a shock wave is produced by a disturbance verynear the surface in the region where laminar (viscous) flow stops andextreme shearing, tearing or ripping takes place. Strong sound waves aregenerated in this region and the relative velocity between the fluid andthe surface is such that the waves cannot escape. A barrier develops andcauses a pressure gradient or shock wave.

Reference is made to FIG. 14 which shows schematically what can beexpected to take place in a binary flow system where steam flows next tothe surface and where diffusion and mixing has not occurred. Inactuality diffusion and mixing does occur and there is no clearcutdistinction between the steam and gas, but for analysis sake, thisgradual change from steam to gas is ignored. The disturbance (tearingaction that generates the sound) takes place within the steam layer andat steam Mach 1 the shock wave starts to develop. As can be noted by thebottom sound wave circles, a ninety degree shock is created such asexists at the throat of any convergent/divergent nozzle. There is noloss in pressure or entropy and no change in temperature at Mach 1 as iswell established. The sound wave in steam travels about 22 percentfaster than in gas and, therefore, as can be noted, the circles flattenout as the waves pass through the gas.

A secondary weaker disturbance can develop where the steam and gas meetwhen theoretically no mixing or diffusion has taken place. The secondarydisturbance will then set up a mild shock wave, the angle of which isdependent upon the gas Mach number. In FIG. 14 a gas Mach number of 1.15is calculated to produce a sixty degree shock wave. Note that the soundwave circles travel some 22 percent slower in the gas and are thussmaller in diameter. Also the sound wave circles bulge out, as can beseen from the dashed projections, when they strike the layer of steam.This mnemonic diagram helps explain the binary flow system and its shockwave (choking) condition which is different in nature than a shock wavefor a monoral flow system.

The nozzle inner and outer side walls 56 and 58 do not present a shockwave problem as their surfaces are smooth and can be shaped to act asany convergent/divergent nozzle where the velocity will increase throughMach 1 without a shock wave and without a loss in entropy.

A further point to consider is that the gas stream exiting at over Mach1 will be diverging due to its change in specific volume relative to itsvelocity whereas the steam layer at Mach less than one will beconverging due to its volume/velocity relationship. Therefore, the totalbinary flow will accelerate at nearly constant cross-sectional areawhich is consistent with there being no physical change in actual areadownstream of the vanes.

If the steam layer is at 1800° F. and at Mach 1, and if the gas streamis at 2200° F., the nozzle exit temperature, then the gas stream canhave a velocity of some 2750 ft/sec. with an equivalent Mach number of1.15. Yet, the main gas stream is independent of the steam layer flow interms of Mach number velocity and only reacts to the sound waves thatoriginate from the primary disturbance or the weaker secondarydisturbance.

The binary flow shock wave phenomenon then leads to the possibility ofbeing able to reach an equivalent Mach 1.15 velocity before full chokingsets in by the steam layer and its shock wave. This being the case, a 15percent higher exit velocity can be achieved without a high trailingedge shock loss occurring. It is important to note that the gastemperature striking the rotating blades will be 100° F. lower becauseof the greater expansion and lower exit pressure. The turbine rotatingblade design and rotating speed can be adjusted upward to accept thishigher velocity and to take advantage of the lower gas temperature. Ahigher work load for this first stage can be achieved without resortingto a transonic reaction rotating blade design and higher transonicfriction and shock wave losses.

Reference is now made to FIG. 15 which shows the normal steam Mach 1throat shock wave 172 in a binary flow system. Steam from the leadingedge weep holes and nozzle number 1 supplies the blanket of steam overthe suction side surface to nozzle 3. Here a trailing edge disturbancetakes place which causes the shock wave to establish itself between thevane trailing edge 34 and nozzle 3 of adjacent vanes as shown. Thisshock wave is just upstream of the Mach 1 nozzle 3 steam exit (800° F.and 2100 ft/sec.). The gas, after leaving the trailing edge, divergesslightly to help fill the trailing edge void and to accelerate to Mach1.15 as it flows from nozzle 3 to the suction side trailing edge. Thevelocity on the pressure side is subsonic all the way from the leadingedge to the trailing edge and only attains Mach 1 at the trailing edgewhere the shock wave takes place. The gas, after leaving the vane, canaccelerate to Mach 1.2 through divergence made possible by the thickerthan normal trailing edge. Finally, the binary flow system of steam andgas offers a way to achieve Mach 1.15 to 1.20 equivalent main streamvelocity without excessive shock wave losses due to the fact that steamhas some 22 percent greater sonic velocity for a given temperature thanthe gas.

Referring further to FIG. 15, there is shown the integrated exit nozzlewidth. The individual steam nozzles 1, 2, 3 and 4 are shown dischargingto the main gas stream. The critical Mach 1 throat width can be seen tobe the arithmetic sum of the three components X, Y and Z. Distance X isshown as the distance from the extreme pressure side vane trailing edgeto the point immediately upstream of steam nozzle 3. Distance Y is shownas the width of steam nozzle 4 exit. Distance Z is shown as the width ofsteam nozzle 3 exit. The total nozzle area can be determined byintegrating the total X, Y and Z distances between the individual vanesfor the 360 degree annulus with the inner and outer diameters of theannulus.

While several embodiments of the nozzle vane assembly and combustorassembly of this invention have been depicted and described, it will beappreciated by those skilled in the art that many variations andmodifications may be made hereto without departing from the fundamentaltheme of the invention.

What is claimed is:
 1. An annular combustor liner for forming acombustion gas comprising; means to direct compressed air into theannular space of said combustor liner, means to direct a combustiblefuel and mix with said compressed air in the annular space of saidcombustor liner, means to ignite said combustible fuel with saidcompressed air and form a combustion gas, said annular combustor linerbeing formed by a plurality of U-shaped elements placed adjacent to eachother, the bend of said U-shaped elements forming the dome of thecombustor liner and the respective legs of said elements forming theinner and outer surfaces of said combustor liner and confining saidannular space, said elements being placed askew to the linear axis ofsaid combustor liner.
 2. The combustor liner of claim 1 including gasturning vanes located at the dome of said combustor liner.
 3. Thecombustor liner of claim 1 including means to direct said fuel into saidcombustor liner in a direction askew to the linear axis of saidcombustor liner.
 4. The combustor liner of claim 1 wherein saidcombustor liner comprises a plurality of independent sections placedadjacent to each other, each section containing a plurality of saidadjacent U-shaped elements.
 5. The combustor liner of claim 1 whereinsaid elements are hollow tubes.
 6. The combustor liner of claim 5wherein said hollow tubes are cladded with an exterior sheathing tocontrol air flow into the annular combustion region.
 7. The combustorliner of claim 1 including a first group of plural spaced openings inthe inner and outer surfaces of said annular combustor between saidU-shaped elements at a location near said fuel mixing means to providethe desired amount of air for controlling combustion within thecombustion zone.
 8. The combustor liner of claim 7 including a secondset of plural spaced openings in the inner and outer surfaces of saidannular combustor liner between said individual spaced elements in alocation downstream of said first group of openings.
 9. The combustorliner of claim 6 including gas turning vanes in the dome of saidcombustor liner.
 10. The combustor liner of claim 9 wherein said gasturning vanes are formed by said sheathing.
 11. The combustor liner ofclaim 10 wherein said fuel mixing means is directed in said annularspace in a flow path askew of said linear axis of said combustor liner.12. The combustor liner of claim 1 wherein the ends of said U-shapedelements forming said outer combustor liner surface are associated witha combustor flange, said flange being bolted to the bulkhead to securesaid combustor liner.
 13. The combustor liner of claim 5 comprising asource of steam and means to direct said steam into the interior of saidtubes.
 14. The combustor liner of claim 4 wherein said independentsections are joined together by matching male and female connectorbracket means.
 15. The combustor liner of claim 4 comprising three ofsaid independent sections.
 16. The combustor liner of claim 1 includingin the dome area a first set of fuel nozzles and an opening for primarycombustion air thereto.
 17. The combustor liner of claim 16 including ata location downstream of said dome area a second set of fuel nozzles andan opening for primary combustion air thereto.
 18. The combustor linerof claim 6 including a source of steam and means to direct steam intothe interior of said tubes and wherein said sheathing is provided withextended surface area to transfer radiant heat to the steam passingthrough the tubes and to provide a means of film cooling the insidesurface of the combustor liner.
 19. The combustor liner of claim 4wherein said independent sections are assembled around a rotatableshaft.
 20. The combustor liner of claim 18 including a narrow spacebetween the sheathing of adjacent tubes to control the amount of coolingair entering the interior of the combustor.
 21. The combustor liner ofclaim 13 including means to direct said steam in the interior of saidtubes from said combustor liner.